Gas turbine engine with modular cores and propulsion unit

ABSTRACT

A separate propulsion unit incorporating a free turbine and a fan receives gases from a plurality of core engines. The core engines each include a compressor, a turbine and a combustion section. The core engines in combination pass gases across the free turbine. A method is also disclosed.

BACKGROUND OF THE INVENTION

This application relates to a gas turbine engine, wherein a core engineis mounted separately from a propulsion unit.

Gas turbine engines are known, and have typically included a fandelivering a portion of air into a bypass duct, and a second portion ofair into a core flow leading into a compressor section. The air iscompressed in the compressor and delivered downstream into a combustionsection where it is mixed with fuel and ignited. Products of thiscombustion pass across turbine rotors which are driven to rotate, and inturn rotate the compressor and fan section. Historically one turbinesection drove both a compressor stage and a fan at the same speed. Morerecently it has been proposed to incorporate a gear reduction such asthe fan can rotate at slower speeds than the compressor stage. With thisarrangement, the outer diameter of the fan can increase, and the outerdiameter of the turbine and compressor sections can decrease.

Historically, the fan and compressors have been mounted coaxially, andhave been driven by turbines that are at a rear end of the engine, withthe fan and compressor at a forward end. It has typically not beenpossible to service any portion of the engine, without removing theconcentrically rotating turbines, compressors and fan as a combinedunit. At a minimum, service is made complex by the inter-relationshipsof these sections.

SUMMARY OF THE INVENTION

In a featured embodiment, a gas turbine engine has a propulsion unitincluding a fan, and a free turbine connected to drive the fan about afirst axis. A plurality of core engines include at least a compressor, acombustion section, and a turbine. The core engine turbine is connectedto drive the compressor. The compressor and core engine turbine rotateabout a second axis. The plurality of core engines has an output fromthe core engine turbines passing over the free turbine.

In another embodiment according to the previous embodiment, a gearreduction is provided between the free turbine and fan.

In another embodiment according to the previous embodiment, the fandelivers propulsion air into a main duct downstream of the fan.

In another embodiment according to the previous embodiment, an inlet tothe core engine is from an ambient inlet separate from the main duct. \

In another embodiment according to the previous embodiment, an inletduct delivers air from the main duct downstream of the fan across theplurality of core engines.

In another embodiment according to the previous embodiment, a connectingduct communicates gases downstream of the core engine turbines from theplurality of core engines across the free turbine.

In another embodiment according to the previous embodiment, gasesdownstream of the common free turbine are directed back into the mainduct.

In another embodiment according to the previous embodiment, gasesdirected downstream of the common free turbine are directed into themain duct through struts.

In another embodiment according to the previous embodiment, core enginesare mounted to rotate about an angle which is generally perpendicular toa rotational axis of the fan.

In another embodiment according to the previous embodiment, the fandelivers propulsion air into a main duct downstream of the fan.

In another embodiment according to the previous embodiment, an inlet tothe core engine is from an ambient inlet separate from the main duct.

In another embodiment according to the previous embodiment, an inletduct delivers air from the main duct downstream of the fan across theplurality of core engines.

In another embodiment according to the previous embodiment, gasesdownstream of the common free turbine are directed back into the mainduct.

In another embodiment according to the previous embodiment, gasesdirected downstream of the common free turbine are directed into themain duct through struts.

In another featured embodiment, a method of providing a gas turbineengine includes the steps of providing a propulsion unit incorporating afree turbine and a fan, and mounting a plurality of core engines to thepropulsion unit, with the core engines each including a compressor, aturbine and a combustor, such that the plurality of core engines incombination provide gases to drive the free turbine.

In another embodiment according to the previous embodiment, free turbinedrives the fan through a gear reduction.

In another embodiment according to the previous embodiment, an inletduct taps air from a main duct downstream of the fan, with the singleinlet duct delivering air into the plurality of core engines.

In another embodiment according to the previous embodiment, the fandelivers propulsion air into a main duct.

In another embodiment according to the previous embodiment, airdelivered into the plurality of core engines coming from an ambient raminlet.

In another embodiment according to the previous embodiment, theplurality of core engines receive inlet air from a tap into the mainduct.

In another embodiment according to the previous embodiment, a singleconnecting duct communicates gases downstream of the plurality of coreengines across the free turbine.

In another featured embodiment, an aircraft incorporating at least onegas turbine engine has an aircraft wing mounting a gas turbine engine.The gas turbine engine includes a propulsion unit including a fan, and afree turbine connected to drive the fan about a first axis, and aplurality of core engines. The core engines include at least acompressor, a combustion section, and a turbine. The core engine turbineis connected to drive the compressor. The compressor and core engineturbine rotate about a second axis. The plurality of core engines has anoutput from the core engine turbines passing over the free turbine.

These and other features of this application will be best understoodfrom the following specification and drawings, the following of which isa brief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically shows a prior art gas turbine engine.

FIG. 2 is a cross-sectional view of an inventive gas turbine engine.

FIG. 3 is a partial view of a portion of the FIG. 2 engine.

FIG. 4 is an alternative embodiment.

FIG. 5A shows a first embodiment.

FIG. 5B shows an alternative embodiment.

FIG. 5C shows yet another embodiment.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a known gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flowpath whilethe compressor section 24 drives air along a core flowpath forcompression and communication into the combustor section 26 thenexpansion through the turbine section 28. Although depicted as aturbofan gas turbine engine in the disclosed non-limiting embodiment, itshould be understood that the concepts described herein are not limitedto use with turbofans as the teachings may be applied to other types ofturbine engines including three-spool architectures.

The engine 20 generally includes a low speed spool 30 and a high speedspool 32 mounted for rotation about an engine central longitudinal axisC relative to an engine static structure 36.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through ageared architecture 48 to drive the fan 42 at a lower speed than the lowspeed spool 30. The high speed spool 32 includes an outer shaft 50 thatinterconnects a high pressure compressor 52 and high pressure turbine54. The inner shaft 40 and the outer shaft 50 are concentric and rotateabout the engine central longitudinal axis C which is collinear withtheir longitudinal axis.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The turbines 46, 54 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion.

An aircraft wing 352 is shown with the gas turbine engine 20 mountedsomewhat forwardly of the engine. A pylon 351 mounts the gas turbineengine to the wing 352. As shown, a disk burst zone A extends for about30° across an exit point of the gas turbine engine. This is an areawhere portions of the gas turbine engine which may fracture, such asportions of the rotor disks, could fly outwardly and damage the wing, asan example. A limitation on the design of where to mount a gas turbineengine is that the disk burse zone A cannot extend across the area wherea fuel tank 400, shown schematically, is mounted. Thus, this hassomewhat limited the mounting of gas turbine engines in the past.

As can be appreciated from FIG. 1, the turbines, compressors are allinter-related and rotate on a common axis with the extending spools30/32. As can be appreciated from FIG. 1, it is somewhat difficult toremove the turbine, compressors, or fans separately from one another forservice.

FIG. 2 shows an inventive engine 100. Air at 114 approaches a fan rotor111 which is driven to rotate with a fan hub 110. A shaft 120 is driventhrough a gear reduction 118, which is in turn driven by a shaft 125.Shaft 125 is driven by a free turbine 127. A duct 310 communicatesproducts of combustion from a core engine 130 that includes low pressureturbine 170, a high pressure turbine 160, a combustor 155, and acompressor section including a high pressure compressor 150 and a lowpressure compressor 145. A spool 165 rotates the low pressure spoolwhile a spool 175 rotates the high pressure spool.

An inlet duct 195 communicates air from a turning duct 185 into the lowpressure compressor 145. An opening 190 takes air from a main duct 105.A pylon 200 mounts the engine 100 to an aircraft wing 352.

A centerline X of core engine 130 incorporating the compressor sections145, 150, combustor 155 and compressor sections 160 and 170 is offset byan angle B from a center line C of the shaft 120/125. Thus, the fanrotor 111 rotates about axis C while the core engine 130 rotates aboutan axis X, which is offset by an angle B. The angle B may be somenon-zero angle, or as described below, may be zero in at least someembodiments. In embodiments which position the core engine to be offset,the angle B may be greater than zero and less than or equal to about90°. Note, other angles can be utilized. The burst zone features aremaximized across this range.

For purposes of the FIG. 2 embodiment, and for moving the burst zone A,the angle B should be greater than zero.

As further shown, a strut 210 extends from the pylon 200 and mounts tothe duct 310.

In the engine 100, rather than delivering air into a core airflow at afan side of the engine, all of the air is delivered into the duct 105. Apropulsion unit including the free turbine 127, gear reduction 118, andfan rotor 111 deliver this air beyond struts 116, and to an outlet 410of a cowl 411. This provides the bulk of the propulsion for the engine.The inlet 190 into the turning duct 185 takes a portion of the air anddelivers it into the inlet 195 for the compressor 145. The air iscompressed, delivered into the higher compressor section 150, into thecombustion section 155, and across turbines 160 and 170, which in turndrive the compressors 150 and 145. Outlet gases downstream of theturbine section 170 passes through the duct 310, and is driven acrossthe free turbine 127. The free turbine 127 drives gear reduction 118 toin turn cause the fan blades 111 to rotate.

Air downstream of the free turbine section 127 passes back outwardly andinto the duct 105 through openings in struts 116.

As can be appreciated from FIG. 2, since the core engine 130 is mountedat an axis which is non-parallel to the axis C, the disk burst zone A isshifted, or angled, forwardly away from the wing 352. Now, the enginemay be mounted further rearwardly underneath the wing than has been thecase in the prior art. Essentially, a core engine, mounted at an axiswhich is non-parallel to the axis of a propulsion unit C would achievethis benefit whenever the axis X is mounted to extend toward the wing352. That is, if the angle B has at least a component extending towardthe wing 352 from the propulsion unit drive axis C, then this forwardmovement of the disk burst zone A will be achieved. The amount ofmovement can be controlled by changing the size of the angle B. A methodof selecting the angle B to position to disk burst zone A such that theengine can be mounted further rearwardly under the wing would also beapparent from the above disclosure.

As can be appreciated in FIG. 3, there are a plurality of struts 116delivering air back into the duct 105. Generally the struts whichdeliver air into the duct are not aligned with the opening 190 into theturning duct 185.

An embodiment 600 is shown schematically in FIG. 4. As shown, a coreengine 608 may communicate gas flow from an inlet duct 606, through acompressor and turbine section as shown in FIG. 2. Products of thecombustion downstream of the turbine sections in the core engine 608pass into a connecting duct 610, and then across a free turbine 612. Thefree turbine 612 may drive the fan rotor 602. The outlet gas from thefree turbine 612 may be directed through the struts 614 and into a mainduct 604. As shown in this Figure, there is a separate propulsion unitincluding the free turbine 612 and fan rotor 602. This may also includea gear reduction in some embodiments. The separate propulsion unit ispositioned forward or toward the inlet of the gas turbine engine 600,while the core engine is spaced rearwardly of the propulsion unit, andis separate from the propulsion unit. With this embodiment, servicing ofthe core engine relative to the propulsion unit is simplified comparedto the prior art.

The fan 602 is positioned at an inlet end of a main air duct 604. Thefree turbine is between the inlet end and the core engine 608 relativeto an axial dimension extending along a rotational axis of the fan, andfrom the inlet end toward an outlet end of the main duct.

FIGS. 5A-5C show the power of having a separate propulsion unit and coreengines in providing modular sizing. The propulsion units 508 can begreatly increased in size. Thus, for example, FIG. 5A shows a 30,000thrust pound engine, FIG. 5B a 60,000 thrust pound engine and FIG. 5C a90,000 thrust pound engine. The propulsion units 508 are all sized upaccordingly. A duct 506 receives ram ambient air from an ambient inletas shown at R, and directs it into a core engine 504. The core engine504 has an output which communicates with a connecting duct 502, whichcommunicates products of combustion back to a free turbine within thepropulsion unit 508. Of course, the use of a turning duct within themain duct, as shown for example in FIG. 2, may be used here.

As shown in the FIG. 5A, the axis of rotation C of the propulsion unitis perpendicular to the axis of rotation X of the core engine 504.

However, other orientations can be utilized such as shown in the aboveembodiments. When it is desired to size the propulsion unit up, a secondcore 504 may be utilized as shown in FIG. 5B. In this manner, only asingle core need to engineered, and several sizes of engine can beachieved by simply adding additional cores. As can be appreciated inFIG. 5B, the size of the ducts 506 and 502 may be larger given the twincores.

FIG. 5C shows the use of three cores. Of course, any number of cores canbe utilized to achieve increased sizing of the thrust capacity of anengine. The internal mechanical and fluid details of the FIG. 5A-Cembodiments may be as shown in FIGS. 1-4.

Details of FIGS. 1-4 are claimed in co-pending U.S. Application Ser. No.13/370,750 filed on even date herewith, now U.S. Pat. No. 8,789,354,granted Jul. 29, 2014, and entitled “Gas Turbine Engine With SeparateCore and Propulsion Unit.”

The core engines not only allow economies from the modular engines, butalso provide redundancies to protect against the failure of any one ofthe core engines.

Although an embodiment of this invention has been disclosed, a worker ofordinary skill in the art would recognize that certain modificationswould come within the scope of this invention. For that reason, thefollowing claims should be studied to determine the true scope andcontent of this invention.

The invention claimed is:
 1. A gas turbine engine comprising: apropulsion unit including a fan, and a free turbine connected to drivesaid fan about a first axis; a plurality of core engines, said coreengines including at least a compressor, a combustion section, and aturbine, said core engine turbine connected to drive said compressor,and said compressor and said core engine turbine rotating about axisnon-coaxial to said first axis; said plurality of core engines having anoutput from said core engine turbines passing over said free turbine;and said propulsion unit being positioned forwardly, and toward andinlet of the gas turbine engine relative to said core engines, and saidcore engines spaced rearwardly of the propulsion unit, and beingseparate from said propulsion unit.
 2. The gas turbine engine as setforth in claim 1, wherein a gear reduction is provided between said freeturbine and said fan.
 3. The gas turbine engine as set forth in claim 1,wherein said fan delivering propulsion air into a main duct downstreamof said fan.
 4. The gas turbine engine as set forth in claim 3, whereinan inlet to said core engines is from an ambient inlet separate fromsaid main duct.
 5. The gas turbine engine as set forth in claim 3,wherein an inlet duct delivers air from said main duct downstream ofsaid fan across said plurality of core engines.
 6. The gas turbineengine as set forth in claim 3, wherein a connecting duct communicatesgases downstream of said core engine turbines from said plurality ofcore engines across said free turbine.
 7. The gas turbine engine as setforth in claim 6, wherein gases downstream of said free turbine aredirected back into said main duct.
 8. The gas turbine engine as setforth in claim 7, wherein gases directed downstream of said common freeturbine are directed into said main duct through struts.
 9. A gasturbine engine comprising: a propulsion unit including a fan, and a freeturbine connected to drive said fan about a first axis; a plurality ofcore engines, said core engines including at least a compressor, acombustion section, and a turbine, said core engine turbine connected todrive said compressor, and said compressor and said core engine turbinerotating about axis non-coaxial to said first axis; said plurality ofcore engines having an output from said core engine turbines passingover said free turbine; and wherein said core engines are mounted torotate about axis which are generally perpendicular to said first axisof said fan.
 10. The gas turbine engine as set forth in claim 9, whereinsaid fan delivering propulsion air into a main duct downstream of saidfan.
 11. The gas turbine engine as set forth in claim 10, wherein aninlet to said core engines is from an ambient inlet separate from saidmain duct.
 12. The gas turbine engine as set forth in claim 10, whereinan inlet duct delivers air from said main duct downstream of said fanacross said plurality of core engines.
 13. The gas turbine engine as setforth in claim 12, wherein gases downstream of said common free turbineare directed back into said main duct.
 14. The gas turbine engine as setforth in claim 13, wherein gases directed downstream of said common freeturbine are directed into said main duct through struts.
 15. An aircraftincorporating at least one gas turbine engine comprising: an aircraftwing mounting a gas turbine engine; said gas turbine engine including apropulsion unit including a fan, and a free turbine connected to drivesaid fan about a first axis, a plurality of core engines, said coreengines including at least a compressor, a combustion section, and aturbine, said core engine turbines connected to drive said compressor,and said compressor and said core engine turbine rotating about axisnon-coaxial to said first axis, said plurality of core engines having anoutput from said core engine turbines passing over said free turbine;and said propulsion unit being positioned forwardly, and toward an inletof the gas turbine engine relative to said core engines, and said coreengines spaced rearwardly of the propulsion unit, and being separatefrom said propulsion unit.